Micro-grooved heat transfer wall

ABSTRACT

A gas turbine engine hot section component such as a turbine blade or vane having an airfoil is provided a non-film cooled portion of a heat transfer wall having a hot surface and a plurality of longitudinally extending micro-grooves disposed in the portion of the wall along the hot surface in a direction parallel to the direction of the hot gas flow. The depth of the micro-grooves is very small and on the order of magnitude of a predetermined laminar sublayer of a turbulent boundary layer. The grooves are sized so as to alter the boundary layer thickness near the leading edge of the airfoil to reduce heat transfer from the hot gas flow to the airfoil near the leading edge. In one embodiment the micro-grooves are about 0.001 inches deep and have a preferred depth range of from about 0.001 inches to 0.005 inches and which are square, rectangular, or triangular in cross-section and the micro-grooves are spaced about one width apart.

RELATED PATENT APPLICATIONS

The present Application deals with related subject matter in co-pendingU.S. patent application Ser. No. 07/733,892, entitled "FILM COOLING OFJET ENGINE COMPONENTS" by Ching-Pang Lee, et al , filed Jul. 22, 1991,assigned to the present Assignee, having two inventors in common withthe present application.

The present Application deals with related subject matter in co-pendingU.S. patent application Ser. No. 08/042,953, entitled "MICRO-GROOVEDHEAT TRANSFER COMBUSTOR WALL" by Steven D Ward, assigned to the presentAssignee.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to cooling of components having non-film cooledhot surfaces for disposal and use in a hot flowpath, and moreparticularly, to components used in gas turbine engines including hotcompressor blades and turbine vanes and blades having non-film cooledhot surfaces.

2. Description of Related Art

It is well known to cool parts using heat transfer across walls havinghot and cold surfaces by flowing a cooling fluid in contact with thecold surface to remove the heat transferred across from the hot surface.Among the various cooling techniques presently used are convection,impingement, and film cooling. These cooling techniques have been usedto cool gas turbine engine hot section components such as turbine vanesand blades. Film cooling has been shown to be very effective butrequires a great deal of fluid flow which typically rewires the use ofpower and is therefore generally looked upon as fuel efficiency andpower penalty in the gas turbine industry.

When turbine components are exposed to a high temperature gas flow, thesurface heat transfer coefficient is inversely proportional to theboundary layer thickness. On airfoil surfaces, the boundary layer growsfrom the leading edge stagnation point toward the trailing edge.Therefore, the boundary layer thickness is small and the heat transfercoefficient is high near the leading edge of the airfoil. Film coolingholes at the leading edge have often been used to cool the leading edgeof the airfoil. It is highly desirable to have a means for altering theboundary layer thickness and reduce the heat transfer coefficient nearthe leading edge.

Another drawback to film cooling is the degree of complexity infabricating and machining the components. In the past, film coolingtechniques have been developed. Turbine airfoils on both blades andvanes often incorporate film cooling holes and slots to flow cooling airalong the hot surfaces of the airfoil walls. Film cooling slots andangled holes require a great deal of fabrication and or machining. Thewall structures themselves are weakened by the cooling airflow passagesrequired to flow the cooling air from the cold to the hot surfaces.

The present invention was developed to improve non-film coolingtechniques for gas turbine engine hot section components so as toefficiently cool the components without resorting to film cooling andthe drawbacks discussed above that are associated with such techniques.

SUMMARY OF THE INVENTION

According to the present invention, a longitudinally extending non-filmcooled hot surface of a heat transfer wall for use in a hot gas flowpathhaving a predetermined flowpath direction is provided withlongitudinally extending micro-grooves in the direction of the flowpath.Typically, micro-grooves are about 0.001 inches deep and have apreferred depth range of from about 0.001 inches to 0.005 inches and aresquare in cross-section, though rectangular, triangular, and othershapes are contemplated by the present invention for other particularapplications and embodiments, and the grooves are spaced about one widthapart. One particular embodiment provides a turbine airfoil withmicro-grooves that extend longitudinally in a generally chordwisedirection over the leading edge portion of the hot surface of theairfoil wall and are disposed in a plurality over substantially thetransversely extending radial length or height of the airfoil.

ADVANTAGES

The present invention provides non-film cooled heat transfer wall suchas the wall of a turbine airfoil with longitudinally extendingmicro-grooves on the hot surface of the wall for exposure to a hot gasflow having a generally longitudinal flowpath direction which improvesthe boundary layer insulation of the hot surface. The depth of themicro-grooves is very small and on the order of magnitude of apredetermined laminar sublayer of a turbulent boundary layer which canbe determined using well known empirical, semi-empirical, and analyticaltechniques. The advantage of such a micro-grooved hot surface is that iteliminates the need for expensive power consuming and sometimesundesirable film cooling holes in certain gas turbine engine hot sectioncomponents. The present invention reduces the heat input into the heattransfer wall. The present invention has the advantage of providing aheat transfer surface for a combustor surface that is more effective inreducing NOx emissions and a leading edge of a turbine airfoil that ismore effectively insulated by the boundary layer. Therefore thecomponent having the heat transfer wall requires less cooling air thanwould otherwise be needed to cool the airfoil.

The foregoing, and other features and advantages of the presentinvention, will become more apparent in the light of the followingdescription and accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing aspects and other features of the invention are explainedin the following description, taken in connection with the accompanyingdrawings where:

FIG. 1 is a cross-sectional view of a gas turbine engine turbine havingnon-film air cooled turbine vane and blade airfoil leading edge wallswith micro-grooved hot surfaces in accordance with the presentinvention.

FIG. 2 is an enlarged perspective view of the combustor linerillustrated in FIG. 1.

FIG. 3 is an enlarged perspective view of the turbine vane airfoilillustrated in FIG. 1.

FIG. 4 is a perspective view of a portion of the heat transfer wall inFIG. 3 having a micro-grooved hot surface in accordance with the presentinvention.

FIG. 5 is a cross-sectional view of the micro-grooved wall in FIG. 3with the riblet having a rectangular cross-section in accordance with analternative embodiment of the present invention.

FIG. 6 is a cross-sectional view of the micro-grooved wall in FIG. 3with the riblet having a thin rectangular cross-section in accordancewith an alternative embodiment of the present invention.

FIG. 7 is a cross-sectional view of the micro-grooved wall in FIG. 3with the riblet having a triangular cross-section in accordance with analternative embodiment of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated in FIG. 1 is a gas turbine engine 10 circumferentiallydisposed about an engine centerline 11 and having in serial flowrelationship a fan section indicated by fan blades 12, a low pressurecompressor 14, a high pressure compressor 16, a combustion section 18, ahigh pressure turbine 20, and a low pressure turbine 21. The combustionsection 18, high pressure turbine 20, and low pressure turbine 21 areoften referred to as the hot section of the engine 10. A high pressurerotor shaft 22 connects, in driving relationship, the high pressureturbine 20 to the high pressure compressor 16 and a low pressure rotorshaft 26 drivingly connects the low pressure turbine 21 to the lowpressure compressor 14 and the fan blades 12. Fuel is burned in thecombustion section 18 producing a very hot gas flow 28 which is flowedto the high pressure and low pressure turbines 20 and 21, respectivelyto power the engine 10. Exposed to the hot gas flow 28 are annular innerand outer micro-grooved combustor liners 30B and 30A, respectively inthe combustion section 18 and turbine blades 31, as shown on the highpressure turbine 20, having externally micro-grooved hot surfaces onleading edges (shown in more particularity in FIG. 3) in accordance withthe present invention. Other hot section components may also make use ofthe micro-grooved surfaces of the present invention (i.e. turbine vanessuch as a turbine inlet guide vane 34). Micro-grooves 37 in hot surfaces36 longitudinally extend downstream in the direction of and parallel tothe predetermined direction of the hot gas flow 28 as indicated by itsarrow.

FIG. 2 illustrates the combustion section 18 and inner and outercombustor liners 30B and 30A, respectively having a hot surface 36exposed to the hot gas flow 28. Micro-grooves 37 in hot surfaces 36 formriblets 38 that are illustrated in more detail in FIG. 4. Note that themicro-grooved hot surface 36 is not film cooled as denoted by lack ofany film cooling apertures or passages from the outer liner cold side39A and inner liner cold side 39B of the inner and outer combustorliners 30B and 30A, respectively. Micro-grooves 37 in hot surfaces 36longitudinally extend downstream in the direction of and parallel to thepredetermined direction of the hot gas flow 28 as indicated by itsarrow.

FIG. 3 illustrates the airfoil 32 having an airfoil wall 33 including ahot surface 36 exposed to the hot gas flow 28. The hot surface 36 hasmicro-grooves 37 disposed over a portion of the airfoil 32 covering aleading edge 40 of the airfoil. Micro-grooves 37 in hot surfaces 36longitudinally extend downstream in the direction of and parallel to thepredetermined direction of the hot gas flow 28 as indicated by itsarrow. One particular embodiment provides a turbine airfoil withmicro-grooves 37 disposed radially apart and that extend longitudinallyin a generally chordwise direction S over the leading edge 40 portion ofthe hot surface 36 of the airfoil wall. Micro-grooves 37 in hot surfaces36 form riblets 38 that are illustrated in more detail in FIG. 4. Notethat the micro-grooved hot surface 36 is not film cooled as denoted bylack of any film cooling apertures or passages from a cold side 39 ofthe airfoil wall 33 to its hot surface 36. The present invention doescontemplate the use of film cooling holes 44 downstream of the leadingedge 40.

Illustrated in FIG. 4 is a portion of a heat transfer wall 42representative of the inner and outer micro-grooved combustor liners 30Band 30A, respectively in FIG. 2 and the airfoil wall 33 in FIG. 3. Onemajor purpose of the invention is to inhibit heat transfer from the hotgas flow 28 to the hot surface 36 of the heat transfer wall 42. This isaccomplished by etching a series of very tiny micro-grooves 37, parallelto the direction of hot gas flow 28 along the hot surface 36, in theheat transfer wall 42 along the hot surface 36 thereby forming theriblets 38 between the micro-grooves. The size of the micro-grooves 37is very small and on the order of magnitude of the predeterminedlocalized thickness of the laminar sublayer of the turbulent boundarylayer which can be determined using well known empirical,semi-empirical, and analytical techniques. The square micro-grooves 37and the riblets 38 between the micro-grooves have an optimalmicro-groove width WG equal to a riblet width WR and an optimalmicro-groove depth D (equal to the height of the riblet 38) depending onthe application. Depth D is defined as the distance between an innermostportion 44 of micro-groove 37 and an outermost portion 50 of adjacentones of riblets 38 as measured in a direction which is substantiallyperpendicular to a plane 52 which passes through an adjacent pair ofriblets 38 located on either side of micro-groove 37. Due to theperspective nature of FIG. 4 plane 52 is not shown in FIG. 4 forpurposes of clarity but is shown in the embodiment of the presentinvention illustrated in FIGS. 5-7. Width WG is defined as the greatestdistance between opposing sidewalls 46 and 48 of micro-groove 37 asmeasured in a direction which is substantially perpendicular to thedirection along which depth D is measured and substantially parallel toplane 52. Opposing sidewalls 46 and 48 are substantially parallel to oneanother in the embodiments illustrated in FIGS. 4-6 but are not parallelin the subsequently discussed embodiment illustrated in FIG. 7.Consequently, the distance between sidewalls 46 and 48 is substantiallyconstant in the embodiments illustrated in FIGS. 4-6 but varies in theembodiment illustrated in FIG. 7. For Reynolds numbers typically foundin combustors and turbines of gas turbine engines micro-grooves having adepth D and a micro-groove width WG on an order magnitude of about 0.001inches are preferred with a preferred range of between 0.001 inches to0.005 inches.

A turbulent boundary layer generally contains eddies and vortexes whichdestroy the laminar regularity of the boundary-layer motion.Quasi-laminar motion persists only in a thin layer in the immediatevicinity of the surface. This portion of a generally turbulent boundarylayer is called the laminar sublayer. The region between the laminarsublayer and completely turbulent portion of the boundary layer iscalled the buffer layer. More information on the subject may be found inone book entitled "Principles of Heat Transfer, Third Edition", by FrankKreith, published by Intext Educational Publishers in 1973, and moreparticularly in chapter 6-3 entitled "Boundary-Layer Fundamentals".

The combination of the air in the micro-grooves 37 and the presence ofthe riblets 38 act to suppress interactions in the boundary layer thatwould otherwise form turbulent bursts which is a mechanism that enhancesheat transfer across a flow. The suppression by the present inventioneffectively reduces the convective heat transfer coefficient, and thus,the heat input to the heat transfer wall 42 by a proportional amount.The present invention contemplates the use of this concept for hot gasturbine engine components such as for use in turbine blades and vanes.The micro-grooves 37 can be etched into the heat transfer wall 42 byseveral methods, such as using an eximer laser, chemical etching,electro-chemical machining (ECM), or electro-discharge machining (EDM),depending on the material of the heat transfer wall.

The invention also contemplates micro-groove cross-sections other thanthe square shape shown in FIG. 4. Illustrated in FIGS. 5 through 7 arethree alternative micro-groove 37 cross-sections for use on a portion ofa heat transfer wall 42. FIG. 5 illustrates a rectangular riblet 38 anda rectangular micro-groove 37 having a micro-groove width WG equal tothe riblet width WR and both of which are smaller than, i.e. about half,the micro-groove depth D. FIG. 6 illustrates a thin rectangular riblet38 and a rectangular micro-groove 37 having a micro-groove width WGlarger than the corresponding riblet width WR, i.e. about five times aslarge. FIG. 7 illustrates a triangular riblet 38 and a triangularmicro-groove 37 having a micro-groove width WG between the tips, oroutermost portion 50, of the triangular riblets 38 larger than, i.e.about twice, the micro-groove depth D. Various embodiments of thepresent invention provide a means for altering a boundary layerthickness proximate a leading edge of an airfoil and for enhancing thestructural integrity of the airfoil, wherein the airfoil is incorporatedin a hot section component such as turbine blade 31 or turbine vane 34,with the means for altering and enhancing comprising a portion of wall42 which is devoid of film cooling apertures, wherein the hot surface 36of wall 42 is operable for exposure to the hot gas flow 28. The meansfor altering and enhancing further comprises the plurality oflongitudinally extending micro-grooves 37 etched in the portion of wall42 along hot surface 36 in a direction substantially parallel to thedetermined flowpath of hot gas flow 28. Micro-grooves 37 are sized so asto increase the boundary layer thickness proximate the leading edge ofthe airfoil and thereby reduce heat transfer from the hot gas flow 28 tothe airfoil proximate the leading edge, relative to an otherwise similarairfoil having a wall portion which is devoid of micro-grooves 37, andmicro-grooves 37 improve the structural integrity of the airfoilrelative to an otherwise similar airfoil having a wall portion includingfilm cooling holes proximate the leading edge.

While the preferred and an alternate embodiment of the present inventionhas been described fully in order to explain its principles, it isunderstood that various modifications or alterations may be made to thepreferred embodiment without departing from the scope of the inventionas set forth in the appended claims.

We claim:
 1. An airfoil for use in a gas turbine engine having a hot gasflow, said airfoil comprising:an airfoil wall including a leading edgeof said airfoil; means for altering a boundary layer thickness proximatesaid leading edge of said air foil and for enhancing the structuralintegrity of said airfoil, said means for altering and enhancingcomprising:a portion of said wall which is devoid of film coolingapertures, said portion comprising at least a part of said leading edgeand said portion having a hot surface operable for exposure to the hotgas flow; and a plurality of longitudinally extending grooves etched insaid portion along said hot surface in a first direction substantiallyparallel to a predetermined flowpath direction of the hot gas flow alongsaid hot surface; wherein said grooves are sized so as to increase theboundary layer thickness proximate said leading edge, and thereby reduceheat transfer from the hot gas flow to said airfoil having a wallportion which is devoid of said grooves, said grooves improving thestructural integrity of said airfoil relative to an otherwise similarairfoil having a wall portion including film cooling holes proximate theleading edge.
 2. An airfoil as claimed in claim 1 wherein said groovesform a plurality of longitudinally extending riblets, each of saidriblets being disposed between an adjacent pair of said grooves, whereineach of said riblets has a lateral cross-sectional shape from a group ofshapes, said group of shapes consisting of square, rectangular, andtriangular shapes.
 3. A component as claimed in claim 1 wherein thecomponent is an airfoil having an airfoil wall including a leading edgeof said airfoil and said portion comprises at least a part of saidleading edge.
 4. An airfoil as claimed in claims 1 or 2 wherein saidgrooves have a depth and a width on the order of magnitude of apredetermined thickness of a laminer sublayer of a turbulent boundarylayer along said portion, said depth being defined as the distancebetween an innermost portion of said groove and an outermost portion ofan adjacent one of said riblets as measured in a second direction whichis substantially perpendicular to said second direction and which issubstantially parallel to said plane.
 5. An airfoil as claimed in claims1 or 2 wherein said grooves have a depth and a width in a range ofbetween about 0.001 inches and 0.005 inches, said depth being defined asthe distance between an innermost portion of said groove and anoutermost portion of an adjacent one of said riblets as measured in asecond direction which is substantially perpendicular to a plane passingthrough said outermost portion of an adjacent pair of said riblets, saidwidth being defined as the greatest distance between opposing sidewallsof said groove as measured in a third direction which is substantiallyperpendicular to said second direction and which is substantiallyparallel to said plane.
 6. An airfoil as claimed in claims 1 or 2wherein said grooves have a depth and a width of about 0.001 inches,said depth being defined as the distance between an innermost portion ofsaid groove and an outermost portion of an adjacent one of said ribletsas measured in a second direction which is substantially perpendicularto a plane passing through said outermost portion of an adjacent pair ofsaid riblets, said width being defined as the greatest distance betweenopposing sidewalls of said groove as measured in a third direction whichis substantially perpendicular to said second direction and which issubstantially parallel to said plane.
 7. An airfoil as claimed in claims1 wherein said grooves have a depth and a width in a range of betweenabout 0.001 inches and 0.005 inches, said depth being defined as thedistance between an innermost portion of said groove and an outermostportion of an adjacent one of said riblets as measured in a seconddirection which is substantially perpendicular to a plane passingthrough said outermost portion of an adjacent pair of said riblets, saidwidth being defined as the greatest distance between opposing sidewallsof said groove as measured in a third direction which is substantiallyperpendicular to said second direction and which is substantiallyparallel to said plane, and wherein said grooves are spaced apart alongsaid hot surface in said third direction by said width.